A typical gas turbine engine has a compressor section, a combustion section and a turbine section. The gas turbine engine includes an annular flowpath for conducting working fluid sequentially through the compressor, combustor, and turbine sections. The compressor has an array of rotating blades and non-rotating stator vanes. The rotating blades compress the working fluid and thereby add momentum to the working fluid. The non-rotating stator vanes orient the flow of working fluid for optimum transfer of energy. Fuel is then added to the compressed working fluid in the combustion section. The mixture of fuel and working fluid is burned in a combustion process which adds energy to the working fluid. The hot working fluid is then expanded through the turbine section. The turbine section includes another array of rotating blades and non-rotating stator vanes. The interaction of the working fluid and the turbine blades transfers energy to the turbine blades. The turbine blades are connected to a rotating shaft which is connected to the compressor blades. In this way the energy which is transferred from the working fluid to the turbine blades is used to compress incoming working fluid in the compressor.
The combustion process raises the temperature of the working fluid in direct proportion to the energy added. The temperature of the working fluid, and thereby the amount of energy which can be added, is limited by the temperature characteristics of the materials used in the turbine section. During operation, rotational forces introduce significant stresses on rotating structure within the turbine section. Increases in temperature reduce the allowable stress and degrade the structural integrity of turbine materials. Therefore, structure within the turbine section must be maintained within acceptable temperature levels. This is especially critical for the first stages of the turbine section which encounter working fluid having the highest temperature.
Structure of particular importance in the turbine section is the rotating seals between the inner diameter of the stator vane assembly and a seal runner extending axially between rotor assemblies. Rotating seals minimize the amount of working fluid which bypasses the blades and vanes, and thereby maximizes the interaction between the working fluid and the airfoil portions of the blades and vanes. A typical rotating seal includes a plurality of radially extending knife-edges disposed on the seal runner which engage an annular shroud of abradable material disposed on the radially inner end of the stator vanes. Gaps exist between the knife-edges and shroud. The size of the gaps must be minimized to maintain proper operation of the rotating seals. Varying operating conditions will produce varying amounts of thermal expansion and thereby variable gap sizes. Therefore, control of the temperature adjacent to the rotating seals is necessary to maintain the gap within acceptable limits.
As is well known in the prior art, a method of maintaining the first stages of the turbine section within acceptable temperature levels is to install a cooling system in the turbine stator vanes. One such cooling system comprises means to inject or conduct cooling air into the body of the hollow stator vanes. Typically compressor bleed air is used as a source of cooling air. In this way cooling is provided to the portion of the stator vanes which extends through the flowpath. The cooling fluid is exhausted through the radially inner portion of the stator vane. A seal cavity, disposed radially inward of the stator vanes, receives the flow of cooling air from the stator vanes. The cooling fluid then cools the rotating seals and other structure local to the seal cavity. Rotating flow surfaces produce a circumferentially flowing, annular body of fluid within the cavity. A drawback to all such cooling systems is the reduced efficiency of the turbine engine as a result of the diversion of working fluid from the compressor section.
Cooling systems for stator vanes and seal cavities have been the focus of much gas turbine research and development. A major focus has been on using the cooling fluid within the seal cavity as efficiently as possible, thereby minimizing the amount of cooling fluid required. Minimizing the cooling fluid taken from the compressor section increases the efficiency of the gas turbine engine.
One example is U.S. Pat. No. 4,869,640, issued to Schwarz et al, entitled "Controlled Temperature Rotating Seal". Schwarz et al discloses structure having a plurality of axially extending, overlapping baffles which cooperate to define multiple mixing volumes within the seal cavity. An intermediate volume is disposed between the working fluid and an innermost volume and provides a thermal buffer between the two. The innermost cavity is partially bounded by and provides cooling fluid for the rotating seals. The thermal buffer prevents direct access of the hot working fluid into the innermost volume. In this way the innermost volume and the rotating seals are maintained at a lower temperature than the intermediate volume.
Aerodynamics of the seal cavity is also a concern as local structure may cause windage losses. Windage losses are the result of the interaction between circumferentially non-continuous flow surfaces and the radially rotating annulus of fluid within the seal cavity. Windage losses generate heat and result in a loss of efficiency for the cooling system and the gas turbine engine. U.S. Pat. No. 4,846,628, issued to Antonellis and entitled "Rotor Assembly for a Turbomachine", is an example of structure which reduces windage losses within the seal cavity. Antonellis discloses a sideplate which is releasably secured to a rotor assembly and has a smooth annular flow surface. The smooth annular flow surface reduces discontinuities within the seal cavity and results in reduced windage losses.
The above art notwithstanding, scientists and engineers under the direction of Applicants' Assignee are working to develop efficient cooling systems for the first stages of the turbine section of a gas turbine engine.